This invention relates to a winged rocket vehicle, and to a method of using the same to efficiently and safely transport payloads to orbital, supraorbital (i.e., Earth escape) or suborbital altitudes and velocities. More specifically, the present invention pertains to a rocket-powered, air-deployed, lift- assisted booster vehicle (ALBV) and to a method of launching the same, which dramatically reduce the amount of rocket propellant and related equipment required to achieve final altitude and velocity for a booster of given payload weight and level of propulsion system technology. In fact, with the current state-of- the-art in space launch vehicle technology, the present invention reduces the total weight of the booster by approximately 50% compared to a comparable ground-launched booster for a given payload weight, with a corresponding reduction in launch system cost.
There is a substantial and continuing commercial and government demand for efficient, economical and reliable payload-carrying space launch vehicles and methods. Numerous prior art approaches to launch payloads into space have been undertaken, but to date, all represent significant compromises regarding safety, economy, reliability and operational flexibility considerations.
Conventional ground-launched ballistic (i.e., non-lifting) booster rockets are the most common prior art approach to payload launch. However, such rockets require complex vertical takeoff facilities, including launch pad apparatus, and are subject to severe operational and geographical restrictions necessitated by the hazards of propellants and flight over populated areas.
Moreover, conventional ground-launched boosters suffer from inherent inefficiencies resulting from a compromise of competing design and operational considerations. These inefficiencies necessarily increase the size, complexity and cost of such systems, making them uneconomic or otherwise undesirable for certain applications.
One such set of competing considerations is the compromise between thrust direction losses and drag losses in conventional ground-launched ballistic boosters. In particular, because the final flight attitude for circular and elliptical orbits, as well as most other missions of interest, is horizontal or substantially horizontal, conventional, vertically launched rockets must pitch over from their initial vertical ascent to a near-horizontal ascent to achieve final orbital flight attitude. Achieving orbit requires high velocity and near-horizontal flight. To minimize losses associated with such thrust direction change (i.e., "thrust direction losses"), pitch-over should ideally occur while the vehicle is ascending at a relatively low velocity, resulting in a near-horizontal ascent early in the trajectory. A shallow ascent profile of this nature was utilized, for example, under zero atmosphere conditions (i.e., in vacuum) by the U.S. Apollo Program Lunar Module to achieve lunar orbit after liftoff from the moon's surface.
Structural stress and aerodynamic heating considerations, however, preclude the implementation of this ideal flight path in applications where the vehicle is being launched through an atmosphere. Aerodynamic forces, including drag and lift forces, increase with the parameter .rho.V.sup.2, where .rho. is the atmospheric density and V is vehicle velocity and the product 1/2.rho.V.sup.2 is the dynamic pressure. Accordingly, for a given velocity, greater drag forces are experienced at lower altitudes than at higher altitudes, since .rho. is greater at such lower altitudes. Because V.sup.2 continuously increases as the vehicle accelerates during booster rocket flight, and .rho. continuously decreases as the vehicle ascends, it is desirable to ascend as near to vertical as possible until the dynamic pressure reaches a maximum value, thereby minimizing the peak aerodynamic load on the vehicle. Accordingly, unlike the zero atmosphere ascent of the Lunar Module, aerodynamic load considerations dictate that conventional ground-launched boosters be launched vertically, with most of the pitch-over from the vertical to the final flight attitude occurring only after .rho.V.sup.2 reaches its maximum value. Consequently, pitch-over occurs at a point where V is extremely high (and .rho. is low), reducing aerodynamic load on the vehicle at the expense of substantial excess propellant usage attributable to thrust direction losses.
In addition, because the conventional ballistic booster spends a significant portion of its flight time in a vertical or near-vertical attitude, the force of gravity directly counteracts the vehicle thrust forces, resulting in other losses, commonly referred to as "gravity losses." Although gravity losses are reduced as a vehicle approaches horizontal flight, the aerodynamic load considerations discussed above preclude substantial horizontal flight of the vehicle until after a maximum value of .rho.V.sup.2 is achieved. Consequently, the conventional booster vehicle incurs substantial gravity losses for a significant portion of its ascent trajectory.
Furthermore, booster rocket motor efficiency increases with increasing exhaust nozzle expansion ratio or nozzle exit area. However, ambient atmospheric pressure forces acting upon the rocket motor nozzle exit are reduce net engine thrust as nozzle area increases. This thrust loss, commonly referred to as "atmospheric pressure-induced thrust reduction," necessitates the design of conventional boosters with nozzle exit areas or expansion ratios providing less than peak motor propulsive efficiency in order to reduce atmospheric pressure-induced thrust reduction and to maximize the net thrust in the denser (lower) regions of the atmosphere.
As is apparent from the foregoing, thrust direction losses, drag losses, gravity losses and atmospheric pressure-induced thrust reduction losses involve complex competing considerations resulting in less than optimum booster performance and flight path maneuvering. Such performance and maneuvering trade-offs greatly increase the size, complexity and expense of conventional boosters for a given payload weight.
To overcome these drawbacks, the present invention proposes launching a lifting-ascent booster vehicle from an aircraft at high altitude and velocity. Launching a booster vehicle from a carrier aircraft while in flight provides the substantial additional advantage of adding the trajectory contributions of the aircraft's velocity and altitude (kinetic and potential energy) directly to the ascent energy of the booster. These trajectory contributions are unavailable for ground-launched booster vehicles.
Another disadvantage of ground-launched vehicles is that the angle of inclination of the resultant orbit relative to the equator is constrained by the latitude of the launch location and by range considerations which limit the direction of launch (i.e., the launch path must not cross populated areas). One of the advantages of launching from an aircraft in flight is that the velocity vector of the aircraft can be aligned with the plane of the final, desired orbit. This is accomplished by flying the carrier aircraft to the desired launch location (at any desired latitude, usually over ocean areas) and giving it the desired velocity vector prior to drop. The principal advantage of being able to fly to the desired location and latitude and in the direction of the desired orbit is that the booster vehicle does not have to perform an energy-consuming inclination change maneuver to achieve the desired orbital inclination, which is much less efficient than using a carrier aircraft to effect the same maneuver.
Another advantage of air-launching over ground launching is the ability to fly to a launch site at any location having favorable weather conditions at the time of launch. Ground launches typically are restricted to only a few selected sites due to safety and security considerations and the availability of the required launch facilities, which usually are at fixed locations. Thus, air launches are less likely than ground launches to be delayed or cancelled due to unfavorable weather conditions.
Various configurations of horizontally launched vehicles have been proposed. However, as will be seen, none provide the advantages in design and operation provided by the present invention.
Jackson, et al., in U.S. Pat. No. 4,265,416, disclose one such system wherein a reusable, winged orbital vehicle is assisted in horizontal ground launch from a runway by one or more reusable, turbojet-propelled, winged booster vehicles that are releasably connected to the orbital vehicle for launch. The boosters assist the rocket-powered orbiter in ascending to staging altitude, and are thereafter released to fly back to Earth for horizontal landing and reuse. The reusable nature of this vehicle requires that it be capable of reentering the atmosphere and using wings to fly back to and land on a runway. In this vehicle, both orbiter and booster wings provide lift, which would have the effect of overcoming gravity losses to some extent. The large size and technical complexity of the disclosed launch vehicle, however, would render it enormously expensive to design, develop and test, and impractical and cost-prohibitive for launching small orbital payloads, e.g., less than 1,000 pounds in weight. Moreover, the size and complexity of the structure, due in large part to its reusable configuration, decrease the payload capacity of the vehicle.
Another launch system using a carrier aircraft to launch a winged booster vehicle while in flight has been proposed by Teledyne Brown Engineering. That system comprises an unmanned spaceplane adapted for horizontal launch from atop a conventional aircraft, such as a Boeing-747. This proposed "piggy-back" technique, however, contemplated starting and testing the spaceplane booster engines while the vehicle is still attached atop the carrier aircraft. Such launching methods are extremely hazardous, substantially limiting wide-scale adoption of this approach. In addition, the system also employs a booster vehicle having wings that remain with the vehicle, which would diminish the payload capacity of the vehicle.
Avoiding the hazards associated with launching from atop a carrier aircraft, certain high-speed research aircraft, e.g., the NASA/North American X-15, have been launched from the underside of carrier aircraft. To date, however, actual underside air launches have been limited to relatively low Mach number, suborbital vehicles, and no vehicle capable of orbital flight has been designed which is suitable for air deployment from a carrier aircraft. The X-15 vehicle attains only about 20% of the energy needed to achieve an orbital trajectory. Furthermore, neither the X-15 nor other aircraft-dropped, rocket-propelled vehicles had two, separable stages, the first providing propulsion as well as lift and aerodynamic control of the trajectory and the second providing propulsion and thrust control of the trajectory. In addition, the reusable configuration of the X-15 vehicle required additional complexity to enable it to survive reentry to the atmosphere and to land on horizontal runways.
There are a number of prior art missiles carried by and launched from carrier aircraft while in flight, including air-to-air and air-to-ground missiles. Such missiles, however, are not designed to leave the atmosphere, do not achieve either orbital speed or altitude, and attain only about 5% of the energy needed to achieve an orbital trajectory. Furthermore, in such missiles the wings and other aerodynamic control surfaces are not jettisoned after ascending to beyond the atmosphere.
It is therefore an object of the present invention to establish an efficient method of launching an orbital, supraorbital or suborbital rocket booster vehicle which reduces the adverse consequences of thrust direction losses, drag losses, gravity losses and atmospheric pressure-induced thrust reduction losses.
It is a further object of the present invention to provide a vehicle which can take advantage of the trajectory energy contributions of a carrier aircraft to increase its payload capacity to Earth orbit and other desired trajectories.
It is still a further object of the present invention to provide a vehicle having expendable wings to reduce vehicle cost and complexity and increase payload capacity.
It is still a further object of the present invention to provide a vehicle for economically and reliably injecting both small and large payloads into orbit.
It is a further object of the invention to provide an orbital, supraorbital or suborbital rocket vehicle that does not require vertical takeoff facilities, and is geographically unrestricted as to its mission departure location, launch location and azimuth, and final orbit inclination, thereby avoiding or minimizing concerns about weather, safety, security and the availability of facilities at a fixed location, which greatly affect the time and location of ground launches.
It is still a further object of the invention to provide a rocket vehicle satisfying the above-mentioned objects which is based on 1988 state-of-the-art propulsion, structures and avionics technologies and devices.
Other objects and advantages will be readily apparent from the following description and drawings which illustrate and described preferred embodiments of the present invention and method of using the same.